Method of clearance control for an interdigitated turbine engine

ABSTRACT

The present disclosure is directed to a method of operating an active clearance control system for an interdigitated turbine engine. The method includes flowing air from a compressor section to a rotatable outer shroud of an interdigitated turbine section; and adjusting the flow of air to the outer shroud based at least on an engine condition of the turbine engine. The present disclosure is further directed to a gas turbine engine including a low speed turbine rotor comprising an inner shroud, an outer shroud, and at least one connecting airfoil coupling the inner shroud and the outer shroud. The outer shroud includes a plurality of outer shroud airfoils extended inward along a radial direction. The engine further includes a turbine frame at least partially surrounding the low speed turbine rotor; a seal assembly disposed between the outer shroud of the low speed turbine rotor and the turbine frame; and a fifth manifold coupled to the turbine frame. The fifth manifold provides a pressurized flow of air to the seal assembly.

FIELD

The present subject matter relates generally to gas turbine engine architecture and operation. More particularly, the present subject matter relates to methods of sealing and clearance control for turbine sections.

BACKGROUND

Gas turbine engines generally include a turbine section downstream of a combustion section that is rotatable with a compressor section to rotate and operate the gas turbine engine to generate power, such as propulsive thrust. General gas turbine engine design criteria often include conflicting criteria that must be balanced or compromised, including increasing fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging (i.e. axial and/or radial dimensions of the engine).

Among turbine sections, define a turbine rotor assembly employing an outer diameter drum to connect rotors. Purge or cooling air is generally necessary to provide sealing and clearance control turbine rotors. Lack of methods and structures for control generally result in excessive air leakage around the relatively large outer diameters of the drums employed to connect rotors. The excessive leakages and lack of control generally result in efficiency losses that may significantly remove the performance and efficiency benefits of an outer drum turbine rotor assembly.

As such, there is a need for structures and methods to mitigate excessive air leakage and control clearances between outer diameter rotor and static structures of the turbine section.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in the following description, or may be obvious from the description, or may be learned through practice of the invention.

The present disclosure is directed to a method of operating an active clearance control system for an interdigitated turbine engine. The method includes flowing air from a compressor section to a rotatable outer shroud of an interdigitated turbine section; and adjusting the flow of air to the outer shroud based at least on an engine condition of the turbine engine.

In one embodiment, adjusting the flow of air to the outer shroud is based at least on an exhaust gas temperature at the turbine section and a rotational speed of the outer shroud. In another embodiment, adjusting the flow of air is further based on a rotational speed of a second turbine rotor interdigitated with a first turbine rotor including the outer shroud.

In various embodiments, adjusting the flow of air to the outer shroud further includes determining a desired clearance gap between the outer shroud and a radially outward end of a plurality of second turbine airfoils of the second turbine rotor. In one embodiment, adjusting the flow of air to the outer shroud further includes determining a desired radial gap between an outer diameter seal assembly and the outer shroud; and determining a desired axial gap between the outer diameter seal assembly and the outer shroud. In still another embodiment, adjusting the flow of air to the outer shroud further includes pressurizing a first side of an outer diameter seal assembly, wherein pressurizing the first side provides a flow of air between the outer shroud and the outer diameter seal assembly. In one embodiment, pressurizing the first side of the outer diameter seal assembly comprises pressurizing an aspirating face seal assembly between the outer shroud and a turbine frame.

In one embodiment, the method further includes flowing air from the compressor section to a second turbine rotor interdigitated with a plurality of outer shroud airfoils of the outer shroud. In another embodiment, flowing air to the second turbine rotor includes flowing air to an inner diameter of the second turbine rotor to attenuate radial thermal growth.

In still various embodiments, the method further includes rotating the gas turbine engine to provide a flow and pressure of air through a primary flowpath and a secondary flowpath; directing at least a portion of the air to an aspirating face seal assembly between the outer shroud and a turbine frame; and directing at least a portion of the air along a longitudinal direction of and around a circumference of the outer shroud based on one or more engine operating conditions. In one embodiment, directing at least a portion of air to the second turbine rotor disposed radially inward of the outer shroud, wherein the portion of air provides thermal attenuation at the second turbine rotor to mitigate radial growth of the second turbine rotor.

In still another embodiment, the method includes determining a desired flow or pressure of air at one or more of the outer shroud and second turbine rotor based on one or more of a lookup table, a transfer function, and a curve comprising at least a function of radial dimensions of the outer shroud and the second turbine rotor relative to one or more of a rotational speed of the outer shroud, a rotational speed of the second turbine rotor, an exhaust gas temperature, a material thermal expansion constant relative to the outer shroud, and a material thermal expansion constant relative to the second turbine rotor.

In various embodiments, the method further includes flowing air from the pressure plenum to a first turbine bearing; flowing air between an outer air bearing of the first turbine bearing and a low speed turbine rotor hub; flowing air between an inner air bearing of the first turbine bearing and a high pressure shaft; and flowing air from the first turbine bearing to a pressure regulating valve. In one embodiment, the method still further includes providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub to a second manifold; and providing a second supply of air from between the inner air bearing and the HP shaft to the second manifold.

The present disclosure is further directed to a gas turbine engine including a low speed turbine rotor comprising an inner shroud, an outer shroud, and at least one connecting airfoil coupling the inner shroud and the outer shroud. The outer shroud includes a plurality of outer shroud airfoils extended inward along a radial direction. The engine further includes a turbine frame at least partially surrounding the low speed turbine rotor; a seal assembly disposed between the outer shroud of the low speed turbine rotor and the turbine frame; and a fifth manifold coupled to the turbine frame. The fifth manifold provides a pressurized flow of air to the seal assembly.

In one embodiment of the engine, the turbine frame defines a longitudinally extended static arm extended between the outer shroud of the low speed turbine rotor and an outer diameter of the turbine frame, and wherein the static arm at least partially defines the seal assembly.

In another embodiment of the engine, the outer shroud of the low speed turbine rotor defines a longitudinally extended rotary arm extended between the turbine frame and the outer shroud. The rotary arm at least partially defines the seal assembly.

In yet another embodiment of the engine, the seal assembly defines a labyrinth seal assembly at the turbine frame and the outer shroud of the low speed turbine rotor.

In still yet another embodiment of the engine, the fifth manifold is coupled to a pressure regulating valve in fluid communication with a second manifold providing pressurized air from a first turbine bearing.

In another embodiment, the engine further includes a compressor section in which the fifth manifold is coupled to the compressor section and extended to the turbine frame to provide a pressurized flow of air to the seal assembly.

These and other features, aspects and advantages of the present invention will become better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the invention and, together with the description, serve to explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including the best mode thereof, directed to one of ordinary skill in the art, is set forth in the specification, which makes reference to the appended figures, in which:

FIG. 1 is a schematic cross sectional view of an exemplary gas turbine engine incorporating an exemplary embodiment of a turbine section according to an aspect of the present disclosure;

FIG. 2 is a schematic cross sectional view of an embodiment of a combustion section and turbine section of the engine shown in FIG. 1 including an embodiment of a first turbine bearing supporting the turbine section;

FIG. 3 is a schematic cross sectional view of an embodiment of the engine including a first turbine bearing supporting the turbine section shown in FIGS. 1-2;

FIG. 4 is a schematic cross sectional view of another embodiment of the engine including a first turbine bearing supporting the turbine section shown in FIGS. 1-2;

FIG. 5 is a schematic cross sectional view of yet another embodiment of the engine including a first turbine bearing supporting the turbine section shown in FIGS. 1-2;

FIG. 6 is a schematic cross sectional view of still another embodiment of the engine including a first turbine bearing supporting the turbine section shown in FIGS. 1-2;

FIG. 7 is a schematic flowchart depicting flows through the engine including a first turbine bearing according to various embodiments shown in FIGS. 1-7;

FIG. 8 is an exemplary flowchart outlining a method of operating the gas turbine engine shown and described in regard to FIGS. 1-7;

FIG. 9 is an exemplary flowchart outlining a method of turbine section thermal management for the gas turbine engine shown and described in regard to FIGS. 1-7;

FIG. 10 is an exemplary flowchart outlining another method of turbine section thermal management for the gas turbine engine shown and described in regard to FIGS. 1-7.

FIG. 11 is an exemplary flowchart outlining a method of clearance control for an interdigitated turbine section such as shown and described in regard to FIGS. 1-7;

FIG. 13 is an exemplary embodiment of an aspirating seal assembly disposed between a rotating outer shroud of the low speed turbine rotor and a turbine frame of the turbine section generally provided in FIGS. 1-7;

FIG. 14 is another exemplary embodiment of an aspirating seal assembly disposed between the rotating outer shroud of the low speed turbine rotor and the turbine frame generally provided in FIGS. 1-7; and

FIG. 15 is yet another exemplary embodiment of an aspirating seal assembly disposed at an outer diameter of the turbine section generally provided in FIGS. 1-7.

Repeat use of reference characters in the present specification and drawings is intended to represent the same or analogous features or elements of the present invention.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention, one or more examples of which are illustrated in the drawings. Each example is provided by way of explanation of the invention, not limitation of the invention. In fact, it will be apparent to those skilled in the art that various modifications and variations can be made in the present invention without departing from the scope or spirit of the invention. For instance, features illustrated or described as part of one embodiment can be used with another embodiment to yield a still further embodiment. Thus, it is intended that the present invention covers such modifications and variations as come within the scope of the appended claims and their equivalents.

As used herein, the terms “first”, “second”, and “third” may be used interchangeably to distinguish one component from another and are not intended to signify location or importance of the individual components.

The terms “upstream” and “downstream” refer to the relative direction with respect to fluid flow in a fluid pathway. For example, “upstream” refers to the direction from which the fluid flows, and “downstream” refers to the direction to which the fluid flows.

The terms “low”, “intermediate”, “high”, or their respective comparative degrees (e.g. -er, where applicable) each refer to relative speeds within an engine unless otherwise specified. For example, a “low turbine” or “low speed turbine” defines a rotational speed lower than a “high turbine” or “high speed turbine”. Alternatively, unless otherwise specified, the aforementioned terms may be understood in their superlative degree. For example, a “low turbine” may refer to the lowest rotational speed turbine within a turbine section, and a “high turbine” may refer to the highest rotational speed turbine within the turbine section.

Exemplary embodiments of methods of clearance control for an interdigitated turbine section are generally provided. The embodiments of the gas turbine engine and methods of operation may enable efficient use of cooling air for operation of the turbine section and mitigation of flowpath leakages, thereby improving overall gas turbine efficiency and/or reducing fuel consumption. The embodiments of the gas turbine engine with interdigitated turbine section shown and described herein may further enable additional stages of interdigitation of the low speed turbine rotor and the high speed turbine rotor, up to and including one or more stages of the low speed turbine rotor forward or upstream of the high speed turbine rotor. In various embodiments, the gas turbine engine with interdigitated turbine section may further reduce cooling air consumption, increase engine efficiency, performance, and/or operability, and/or reduce part quantities, weight, and/or packaging (i.e. axial and/or radial dimensions). Still further, the interdigitated turbine section may reduce a product of axial flow area and the square of the rotational speed (the product referred to as “AN²”) while additionally reducing an average work factor per stage of the turbine section.

Referring now to the drawings, FIG. 1 is a schematic cross sectional view of an exemplary gas turbine engine 10 (herein referred to as “engine 10”), shown as a high bypass turbofan engine, incorporating an exemplary embodiment of a turbine section 90 according to an aspect of the present disclosure. Although further described below with reference to a turbofan engine, the present disclosure is also applicable to turbomachinery in general, including propfan, turbojet, turboprop, and turboshaft gas turbine engines, including marine and industrial turbine engines and auxiliary power units. Still further, although described below as a three-spool gas turbine engine, the present disclosure is also applicable to two-spool gas turbine engines. As shown in FIG. 1, the engine 10 has a longitudinal or axial centerline axis 12 that extends there through for reference purposes. The engine 10 defines a longitudinal direction L, a radial direction R, and an upstream end 99 and a downstream end 98 along the longitudinal direction L.

In general, the engine 10 may include a substantially tubular outer casing 18 that defines an annular inlet 20. The outer casing 18 encases or at least partially flows, in serial flow arrangement, a compressor section 21, a combustion section 26, and an interdigitated turbine section 90 (herein referred to as “turbine section 90”). Generally, the engine 10 defines, in serial flow arrangement from the upstream end 99 to the downstream end 98, the fan assembly 14, the compressor section 21, the combustion section 26, and the turbine section 90. In the embodiment shown in FIG. 1, the compressor section 21 defines a high pressure (HP) compressor 24 and an intermediate pressure (IP) compressor 22. In other embodiments, the fan assembly 14 may further include or define one or more stages of a plurality of fan blades 42 that are coupled to and extend outwardly in the radial direction R from a fan rotor 15 and/or a low speed shaft 36. In various embodiments, multiple stages of the plurality of fan blades 42 coupled to the low speed shaft 36 may be referred to as a low pressure (LP) compressor.

An annular fan casing or nacelle 44 circumferentially surrounds at least a portion of the fan assembly 14 and/or at least a portion of the outer casing 18. In one embodiment, the nacelle 44 may be supported relative to the outer casing 18 by a plurality of circumferentially-spaced outlet guide vanes or struts 46. At least a portion of the nacelle 44 may extend over an outer portion (in radial direction R) of the outer casing 18 so as to define a bypass airflow passage 48 therebetween.

Referring now to FIG. 2, an exemplary embodiment of the turbine section 90 of the engine 10 is generally provided. The turbine section 90 includes a low speed turbine rotor 110 extended along the longitudinal direction L. The low speed turbine rotor 110 includes an inner shroud 112, an outer shroud 114, and at least one connecting airfoil 116 coupling the inner shroud 112 to the outer shroud 114. The outer shroud 114 includes a plurality of outer shroud airfoils 118 extended inward along the radial direction R. In various embodiments, the inner shroud 112 may include a plurality of inner shroud airfoils 119 extended outward along the radial direction R.

The inner shroud 112 and the outer shroud 114 each extend generally along the longitudinal direction L. The inner shroud 112 and/or the outer shroud 114 may each extend at least partially in the radial direction R. In various embodiments, the inner shroud 112 extends from the connecting airfoil 116. In one embodiment, the inner shroud 112 further extends toward the downstream end 98 along the longitudinal direction L. In still various embodiments, the outer shroud 114 extends from the connecting airfoil 116 toward the upstream end 99 along the longitudinal direction L toward the combustion section 26.

Between the outer shroud 114 and a turbine frame 295 are generally one or more seal assemblies 399, 400 disposed therebetween to prevent and control purge cavity flow and pressure (e.g., secondary flowpath 71 shown in FIGS. 4-7). In one embodiment, the seal assembly 399 defines a labyrinth seal. For example, the turbine frame 295 may define a plurality of grooves or pathways and the outer shroud 114 may define one or more teeth extending within the groove. The pressure differentials and flow vortices between the grooves and teeth mitigate flow, and therefore leakage, across the seal assembly 399. In other embodiments, the seal assembly 400 defines an aspirating face seal assembly, such as shown and described further in regard to FIGS. 13-15. In various embodiments, the turbine section 90 and engine 10 defines one or more of the seal assemblies 399, 400, or combinations thereof.

As shown on FIG. 4, a turbine casing 155 extends generally extended from the combustion section 26 along the longitudinal direction L and circumferentially encompassing the turbine section 90. The turbine casing 155 may define an outer diameter (OD) secondary flowpath 71 radially between the outer shroud 114 of the low speed turbine rotor 110 and the turbine casing 155. The turbine casing 155 may include a plurality of orifices 156 through which a fluid may ingress or egress to the secondary flowpath 71 and/or core flowpath 70.

Referring now to FIG. 2, the turbine section 90 further includes a high speed turbine rotor 120 and an intermediate speed turbine rotor 130 each disposed forward or upstream 99 of the one or more connecting airfoils 116 of the low speed turbine rotor 110. The high speed turbine rotor 120 includes a plurality of high speed turbine airfoils 122 extended outward along the radial direction R. The intermediate speed turbine rotor 130 includes a plurality of intermediate speed turbine airfoils 132 extended outward along the radial direction R. The pluralities of high speed turbine airfoils 122 and intermediate speed turbine airfoils 132 are each disposed among the pluralities of outer shroud airfoils 118 of the low speed turbine rotor 110 along the longitudinal direction L.

In various embodiments, the low speed turbine rotor 110 defines a plurality of stages of rotating airfoils, such as the plurality of outer shroud airfoils 118 disposed along the longitudinal direction L, the one or more connecting airfoils 116, and/or the plurality of inner shroud airfoils 119 disposed along the longitudinal direction L. In one embodiment, the low speed turbine rotor 110 defines at least one stage forward or upstream 99 of the high speed turbine rotor 120. In another embodiment, the turbine section 90 defines a first stage of airfoils in which the first stage includes the plurality of outer shroud airfoils 118 of the low speed turbine rotor 110 forward or upstream 99 of each stage of the high speed turbine rotor 120.

In still various embodiments, such as shown in FIG. 2, the engine 10 defines, in serial flow arrangement along the longitudinal direction L from the upstream end 99 to the downstream end 98, the plurality of outer shroud airfoils 118 of the low speed turbine rotor 110, the plurality of high speed turbine airfoils 122 of the high speed turbine rotor 120, and the plurality of outer shroud airfoils 118 of the low speed turbine rotor 110. In still various embodiments, additional iterations of interdigitation of the low speed turbine rotor 110 and the high speed turbine rotor 120 may be defined forward or upstream 99 of the connecting airfoils 116.

In yet various embodiments, such as shown in FIG. 2, the engine 10 further defines two or more stages of the low speed turbine rotor 110 and one or more stages of the intermediate speed turbine rotor 130 in interdigitation along the longitudinal direction L downstream of the high speed turbine rotor 120. For example, the engine 10 may define, in serial flow arrangement along the longitudinal direction L from the upstream end 99 to the downstream end 98, the combustion section 26, the first stage 101 of the low speed turbine rotor 110, the high speed turbine rotor 120, the low speed turbine rotor 110, the intermediate speed turbine rotor 130, and the low speed turbine rotor 110. In one embodiment, the engine 10 may further define additional stages of interdigitation of the high speed turbine rotor 120 and/or the intermediate speed turbine rotor 130 in interdigitation (i.e. alternating arrangement along the longitudinal direction L) with the low speed turbine rotor 110.

As another non-limiting example, as shown in FIG. 2, the engine 10 may further define the serial flow arrangement of the plurality of outer shroud airfoils 118, the plurality of high speed turbine airfoils 122, the plurality of outer shroud airfoils 118, the plurality of intermediate speed turbine airfoils 132, the plurality of outer shroud airfoils 118, an additional plurality of intermediate speed turbine airfoils 132, and the connecting airfoils 116. It should be appreciated that although FIG. 2 shows the high speed turbine rotor 120 as defining one stage, the high speed turbine rotor 120 may define generally one or more stages between the first stage 101 of the low speed turbine rotor 110 and the connecting airfoils 116 of the low speed turbine rotor 110, and interdigitated therebetween along the longitudinal direction L. Similarly, it should be appreciated that although FIG. 1 shows the intermediate speed turbine rotor 130 as defining two stages, the intermediate speed turbine rotor 130 may define generally one or more stages between the high speed turbine rotor 120 and the connecting airfoils 116 of the low speed turbine rotor 110.

Referring now to FIGS. 1-7, in various embodiments, the low speed turbine rotor 110 is drivingly connected and rotatable with a low speed shaft 36 extended along the longitudinal direction L and generally concentric about the axial centerline 12. In one embodiment, as shown in FIG. 1, the low speed shaft 36 is connected to the fan assembly 14, of which is driven in rotation by the low speed turbine rotor 110 of the turbine section 90. The low speed shaft 36 is connected to the fan rotor 15 of the fan assembly 14. In various embodiments, the fan assembly 14 may define a plurality of stages of the plurality of fan blades 42, of which further define the LPC.

In various embodiments, the intermediate speed turbine rotor 130 is drivingly connected and rotatable with an intermediate speed shaft 35 extended along the longitudinal direction L and generally concentric about the axial centerline 12. In one embodiment, as shown in FIG. 1, the intermediate speed shaft 35 is connected to the IP compressor 22, of which is driven in rotation by the intermediate speed turbine rotor 130 of the turbine section 90.

Referring still to FIGS. 1-7, the high speed turbine rotor 120 of the turbine section 90 is drivingly connected and rotatable with a HP shaft 34 extended along the longitudinal direction L and generally concentric about the axial centerline 12. The HP shaft 34 is connected to the HP compressor 24, of which is driven in rotation by the high speed turbine rotor 120 of the turbine section 90.

Referring to FIG. 2, the turbine section 90 further includes a static turbine frame 295 extended generally annularly around the axial centerline 12. In various embodiments, the turbine frame 295 defines an outer turbine casing at least partially surrounding the turbine section 90 along the longitudinal direction. In still various embodiments, the turbine frame 295 defines one or more of an inter-turbine frame (e.g., between two or more turbine rotors 110, 120, 130), an aft turbine frame disposed aft of the turbine rotors 110, 120, 130, and a forward turbine frame, such as disposed longitudinally between the combustion section 26 and the turbine rotors 110, 120, 130.

In one embodiment, the turbine frame 295 includes one or more turbine vanes 150. The turbine vane 150 may define a plurality of stationary airfoils (i.e. vanes) in circumferential arrangement. In one embodiment, the turbine vane 150 is disposed between the pluralities of inner shroud airfoils 119 along the longitudinal direction L. In various embodiments, the turbine vane 150 is disposed downstream 98 of the connecting airfoil 116 of the low speed turbine rotor 110. The turbine vane 150, or pluralities thereof, interdigitated among the pluralities of inner shroud airfoils 119 may enable further conditioning of the combustion gases 86 and work or energy extraction from the low speed turbine rotor 110 via the plurality of inner shroud airfoils 119.

During operation of the engine 10 as shown collectively in FIGS. 1-7, the high speed turbine rotor 120 rotates generally at a higher rotational speed than the intermediate speed turbine rotor 130. The intermediate speed turbine rotor 130 rotates generally at a higher speed than the low speed turbine rotor 110. During operation of the engine 10, a volume of air as indicated schematically by arrows 74 enters the engine 10 through an associated inlet 76 of the nacelle and/or fan assembly 14. As the air 74 passes across the fan blades 42, a portion of the air as indicated schematically by arrows 78 is directed or routed into the bypass airflow passage 48 while another portion of the air as indicated schematically by arrows 80 is directed or through the fan assembly 14. Air 80 is progressively compressed as it flows through the compressor section 21 toward the combustion section 26.

The now compressed air, as indicated schematically by arrows 82, flows into the combustion section 26 where a fuel 91 is introduced, mixed with at least a portion of the compressed air 82, and ignited to form combustion gases 86. The combustion gases 86 flow into the turbine section 90, causing rotary members of the turbine section 90 to rotate and support operation of respectively coupled rotary members in the compressor section 21 and/or fan assembly 14.

In various embodiments, the low speed turbine rotor 110, and the low speed shaft 36 to which it is attached, rotates in a first direction along the circumferential direction. The high speed turbine rotor 120, and the HP shaft 34 to which it is attached, rotates in a second direction opposite of the first direction along the circumferential direction. In one embodiment, the intermediate speed turbine rotor 130, and the intermediate speed shaft 35 to which it is attached, rotates in the second direction in co-rotation with the high speed turbine rotor 120 and in counter-rotation with the low speed turbine rotor 110.

It should further be understood that the first direction and the second direction as used and described herein are intended to denote directions relative to one another. Therefore, the first direction may refer to a clockwise rotation (viewed from downstream looking upstream) and the second direction may refer to a counter-clockwise rotation (viewed from downstream looking upstream). Alternatively, the first direction may refer to a counter-clockwise rotation (viewed from downstream looking upstream) and the second direction may refer to a clockwise rotation (viewed from downstream looking upstream).

Still further during an operation of the engine 10, combustion gases 86 exiting the combustion section 26 define a generally low speed toward the downstream end 98 of the engine 10. A low speed rotation (e.g. along a tangential or circumferential direction) of the first stage 101 of the low speed turbine rotor 110 accelerates a speed of the combustion gases 86, such as in the tangential or circumferential direction, to approximately equal or greater than a speed of the high speed turbine rotor 120.

By defining the low speed turbine rotor 110 as the first stage 101 of the turbine section 90 aft or downstream of the combustion section 26, the engine 10 may obviate the need for a first turbine vane or nozzle guide vane to accelerate the combustion gases 86 forward or upstream of the high speed turbine rotor 120 defining a high speed turbine. As such, the engine 10 may reduce a quantity or mass of cooling air from the compressor section 21 and/or combustion section 26, thereby increasing engine efficiency by enabling more energy (i.e. compressed air) to be used during combustion. Additionally, or alternatively, the turbine section 90 may reduce necessary cooling air and enable increased performance and/or operability of the compressor section 21, including surge margin and/or efficiency, or decrease a required amount of work from the compressor section 21, which may reduce axial dimensions or stages of the compressor section 21 and further reduce engine packaging, weight, and/or part count, and generally improve engine 10 performance.

Additionally, obviating the need for the first turbine vane or nozzle guide vane may enable the turbine section 90, or more specifically, the first stage 101, as a rotating stage, to be designed to an average combustion gas 86 temperature rather than designed to accommodate peak temperatures (i.e. high spots) along an annulus of the core flowpath 70 within the combustion section 26. Therefore, as all of the plurality of outer shroud airfoils 118 of the first stage 101 are rotating, all of the plurality of outer shroud airfoils 118 may only transiently endure adverse effects of combustion hot spots rather than substantially sustained or constant exposure to a higher temperature from the combustion gases in contrast to other locations about the annulus of the core flowpath 70. Still further, the turbine section 90 described herein may enable alternative design methods for the combustion section 26 due to a decreased adverse effect of combustion hot spots on the turbine section 90. Therefore, the turbine section 90 may enable design of the combustion section 26 to further improve combustion stability, decrease emissions, increase operability across all or part of a flight envelope, increase altitude re-light performance, and/or decrease lean blowout (LBO).

Referring now to FIG. 3, a detailed schematic cross sectional view of the engine 10 including a first turbine bearing 200 supporting the low speed turbine rotor 110 and the high speed turbine rotor 120 is generally provided. The first turbine bearing 200 is disposed radially between the hub 105 of the low speed turbine rotor 110 and the HP shaft 34 coupled to the high speed turbine rotor 120. The first turbine bearing 200 defines an outer air bearing 210 along an outer diameter 215 of the first turbine bearing 200 and adjacent to the hub 105 of the low speed turbine rotor 110. The first turbine bearing 200 further defines an inner air bearing 220 along an inner diameter 225 and adjacent to the HP shaft 34 coupled to the high speed turbine rotor 120.

In various embodiments, the first turbine bearing 200 includes an outer annular wall 211 at the outer diameter 215 of the outer air bearing 210 and an inner annular wall 221 at the inner diameter 225 of the inner air bearing 220. Each of the outer and inner annular walls 211, 221 extend circumferentially around the axial centerline 12 of the engine 10 and are generally concentric to the axial centerline 12. The outer annular wall 211 and inner annular wall 221 each extend at least partially along the longitudinal direction L. The first turbine bearing 200 further includes an outer plurality of orifices 213 arranged at least along the longitudinal direction L and circumferentially along the outer air bearing 210. The outer plurality of orifices 213 is adjacent to the hub 105 of the low speed turbine rotor 110. The first turbine bearing 200 still further includes an inner plurality of orifices 223 arranged at least along the longitudinal direction L and circumferentially along the inner air bearing 220. The inner plurality of orifices 223 is adjacent to the HP shaft 34 coupled to the high speed turbine rotor 120.

In various embodiments, the outer air bearing 210 and/or the inner air bearing 220 may be formed of a plurality of segments arranged substantially in circumferential arrangement about the axial centerline 12 of the engine 10 to together define an approximately 360 degree annular surface, such as the outer annular wall 211 and/or the inner annular wall 221.

Referring to FIG. 3, the combustion section 26 includes a compressor/diffuser frame 25 generally surrounding a combustor assembly and defining a pressure plenum 84. The compressor/diffuser frame 25 generally defines a pressure vessel defining the annular pressure plenum 84 downstream of an exit of the HP compressor 24 and generally upstream of and/or radially surrounding the combustion chamber 85 at which the compressed air 82 substantially defines a compressor exit pressure. The compressor/diffuser frame 25 generally includes an inner diffuser frame 27 generally annular and radially inward within the combustion section 26. The inner diffuser frame 27 generally extends from the compressor section 21 toward the turbine section 90, supporting the first turbine bearing 200 and a combustor assembly. The first turbine bearing 200 is coupled to and cantilevered from the inner diffuser frame 27 toward the turbine section 90.

In various embodiments, the first turbine bearing 200 defines an at least partially annular groove 217 at the outer diameter 215 of the outer annular wall 211 of the outer air bearing 210. The groove 217 is disposed radially adjacent to the hub 105 of the low speed turbine rotor 110. In various embodiments, the groove 217 is defined proximate to an upstream end of the hub 105 of the low speed turbine rotor 110. For example, the groove 217 may be defined on the outer air bearing 210 radially adjacent to within 50% of the longitudinal span of the hub 105 from the upstream end of the hub 105 (i.e., from the end of the hub 105 most proximate to the upstream end 99 of the engine 10). As another example, the groove 217 may be defined on the outer annular wall 211 of the first turbine bearing 200 adjacent along the radial direction R within 30% of the span of the hub 105 along the longitudinal direction L proximate to the upstream end 99 of the engine 10.

In one embodiment, the engine 10 further defines a compressor seal interface 240 between the HP compressor 24 and the first turbine bearing 200. Between the compressor seal interface 240 and the first turbine bearing 200, the engine 10 defines a first cavity 245 extended at least partially circumferentially around the axial centerline 12 of the engine 10. The engine 10 further defines a turbine seal interface 250 between the high speed turbine rotor 120 and the first turbine bearing. Between the turbine seal interface 250 and the first turbine bearing 200, the engine 10 defines a second cavity 255 extended at least partially circumferentially around the axial centerline 12 of the engine 10. The seal interfaces 240, 250 may each include a combination of a shroud and knife edge or labyrinth seal to define each seal interface 240, 250.

The engine 10 may further include a first manifold 260 extended from the pressure plenum 84 to the first turbine bearing 200 and in fluid communication therebetween. The compressed air 82 from the compressor section 21 may flow from the pressure plenum 84 through the first manifold 260 into the first turbine bearing 200 as shown schematically by arrows 261. The air 261 through the first turbine bearing 200 flows to and egresses through the first and second plurality of orifices 213, 223 at the outer air bearing 210 and the inner air bearing 220 of the first turbine bearing 200.

The first turbine bearing 200 generally defines a hydrostatic air bearing through which the compressed air 261 egresses through the first and second plurality of orifices 213, 223 to define a separating fluid film between the rotating components. At the outer air bearing 210 of the first turbine bearing 200, the compressed air 261 defines a fluid film between the outer annular wall 211 and the hub 105 of the low speed turbine rotor 110. At the inner air bearing 220 of the first turbine bearing 200, the compressed air 261 defines a fluid film between the inner annular wall 221 and the HP shaft 34 coupled to the high speed turbine rotor 120. The pressurized air from the compressor section 21 and in the pressure plenum 84 at least partially defines or determines an amount of clearance or distance between the hub 105 of the low speed turbine rotor 110 and the first turbine bearing 200, and between the HP shaft 34 of the high speed turbine rotor 120 and the first turbine bearing 200.

Referring now to FIG. 4, the first turbine bearing 200 shown and described in regard to FIG. 3 may further define a cooling orifice 230 axially adjacent to the high speed turbine rotor 120 through which a portion of the compressed air 261 within the first turbine bearing 200 may egress into the high speed turbine rotor 120 to provide thermal attenuation or cooling, such as shown schematically by arrows 231. In various embodiments, the high speed turbine rotor 120 defines a turbine cooling circuit 125 extended therethrough at least partially along the longitudinal direction L and radial direction R. The turbine cooling circuit 125 is defined generally within the structure or walls of a rotor, hub, or disk of the high speed turbine rotor 120. The turbine cooling circuit 125 may further be defined within the structure of the high speed turbine airfoils 122. The turbine cooling circuit 125 may define one or more cooling circuit egresses 127 at the high speed turbine airfoils 122, through which the compressed air 231 egresses the high speed turbine rotor 120 into the core flowpath 70 at the turbine section 90.

In various embodiments, the cooling orifice 230 of the first turbine bearing 200 defines a nozzle to increase a speed of and/or direct a flow of air 231 to the cooling circuit 125 of the high speed turbine rotor 120. For example, the cooling orifice 230 may define a plurality of vanes defining a pressure side, a suction side, a leading edge, and a trailing edge, across which the compressed air 231 accelerates a flow of the air 231 along the longitudinal direction L and/or along the circumferential direction to approximately match the rotational speed of the high speed turbine rotor 120. In various embodiments, the plurality of vanes defined within the cooling orifice 230 may further accelerate the flow of air 231 along the same direction of rotation as the high speed turbine rotor 120 (i.e., co-directional or co-rotational). For example, in the embodiments described herein, the plurality of vanes of the cooling orifice 230 may accelerate the air 231 generally along the second direction along the circumferential direction generally co-directional to the high speed turbine rotor 120 and opposite of the low speed turbine rotor 110. However, in various embodiments, the cooling orifice 230 may accelerate the flow of air along the first direction.

Referring now to FIGS. 3-4, as the compressed air 261 pressurizes the space between the outer air bearing 210 and the hub 105, a portion of the compressed air, shown schematically as 262, may egress into the core flowpath 70. Furthermore, as the compressed air 261 pressurizes the space between the first turbine bearing 200 and the hub 105 of the low speed turbine rotor 110, the compressed air 261 is collected in an annular cavity 218 defined by the groove 217 at the outer annular wall 211. Furthermore, as the compressed air 261 pressurizes the space between the first turbine bearing 200 and the HP shaft 34 coupled to the high speed turbine rotor 120, the compressed air 261 is collected in the first cavity 245 defined between the compressor seal interface 240 and the first turbine bearing 200. In various embodiments, the compressed air 261 may further be collected in the second cavity 255 defined between the turbine seal interface 250 and the first turbine bearing 200. The air collected in each cavity 218, 245, 255 may then be collected within the first turbine bearing 200 and egressed through a second manifold 270 as shown schematically by arrows 271. The second manifold 270 is extended from the first turbine bearing 200 to a pressure regulating valve 300 and in fluid communication therebetween.

In various embodiments, the second manifold 270 is in fluid communication with the annular cavity 218 and the first cavity 245. The second manifold 270 may define a flow and pressure of the compressed air 271 therethrough defined by the air from the annular cavity and the first cavity. In another embodiment, the second manifold 270 is further in fluid communication with the second cavity 255. The second manifold 270 may define a flow and pressure of the compressed air 271 therethrough defined by the air from the annular cavity, the first cavity, and the second cavity.

In various embodiments, the engine 10 further includes a third manifold 280 extended from the pressure plenum 84 of the combustion section 26 to the pressure regulating valve 300 and in fluid communication therebetween. In one embodiment, the third manifold 280 communicates a reference pressure or motive pressure to the pressure regulating valve 300. The compressed air 82 in the pressure plenum 84 pressurizes the third manifold 280 as shown schematically by arrows 281.

During operation of the engine 10 shown in FIGS. 1-4, the compressed air 281 from the pressure plenum generally provides a reference pressure to the pressure regulating valve 300 that may generally reflect or correspond to the pressure in pressure plenum 84. The pressure regulating valve 300 generally regulates a difference in pressure between the third manifold 280 and the compressed air 281 within it and the second manifold 270 and the compressed air 271 egressing the first turbine bearing 200. As such, the pressure regulating valve 300 regulates the pressure of air between the hub 105 of the low speed turbine rotor 110 and the outer air bearing 210 as well as the pressure of air between the inner air bearing 220 and the HP shaft 34 coupled to the high speed turbine rotor 120.

Referring now to FIG. 5, the engine 10 may be configured substantially similarly to the embodiments shown and described in regard to FIGS. 1-4. In FIG. 5, the engine 10 may further define an OD turbine cooling supply manifold 275 extended from the second manifold 270 and coupled to the OD secondary flowpath 71 of the turbine section 90. In one embodiment, a one way valve 310 is coupled to the turbine cooling supply manifold 275 to enable a flow of the compressed air 271 to the OD secondary flowpath 71 of the turbine section 90. In various embodiments, the one way valve 310 may be referred to as a check valve or non-return valve, enabling a single direction of flow along the second manifold 270 from the first turbine bearing 200.

During operation of the engine 10 shown in FIG. 5, when the pressure in the pressure plenum 84 and, subsequently, the third manifold 280, is low, the pressure regulating valve 300 opens and sends the compressed air, as shown schematically by arrows 299, to a lower pressure sink. In one embodiment, the lower pressure sink includes the core flowpath 70 toward a downstream end 98 of the turbine section 90. In another embodiment, the lower pressure sink includes an ambient area radially outside of the engine 10 (e.g., at approximately ambient or atmospheric pressure condition, such as at approximately 1 bar, or less at altitudes above sea level). When the pressure regulating valve 300 opens and sends the compressed air 299 to the lower pressure sink, the one way valve 310 generally closes. Conversely, when pressure in the pressure plenum 84 is high, the one way valve 310 generally opens and enables a flow of the compressed air, as shown schematically by arrows 276, into the turbine section 90. The compressed air 276 provides cooling to the interdigitated low speed turbine rotor 110, such as the outer shroud 114 and one or more of the plurality of outer shroud airfoils 118 extended therefrom. The compressed air 276 may additionally, or alternatively, provide buffering at one or more seals within the OD secondary flowpath 271 to restrict a flow of air from higher pressure cavities (e.g., from the combustion section 26 or generally from higher pressures upstream of others within the OD secondary flowpath 271 of the turbine section 90).

When included in an aircraft, the engine 10 may define the pressure in the pressure plenum 84 as low or high based, at least in part, on a landing and take-off (LTO) cycle of the aircraft, such as, but not limited to, those defined by standards, regulations, or requirements from the Federal Aviation Administration or the European Aviation Safety Agency. For example, the LTO as known in the art generally includes taxi, take-off, climb, approach, landing, and taxi. The LTO may further include cruise following take-off and descent following cruise and prior to approach. Generally, the engine 10 defines a high pressure at the pressure plenum 84 at a cruise condition or higher, such as including climb and take-off. The engine 10 defines a low pressure at the pressure plenum 84 below a cruise condition, such as at approach, landing, and taxi. Although various embodiments of the engine 10 may define quantitatively different high and low pressures at the pressure plenum 84, it should be understood that various embodiments of the engine 10 may define qualitatively similar pressures relative to the LTO.

It should be appreciated that although the engine 10 may define the aforementioned LTO during normal operation, deviations in the LTO may define a high pressure in the engine 10 at lower altitudes, such as, but not limited to, altitudes commensurate to approach conditions or lower. For example, the engine 10 may define a high pressure approximately similar to the cruise condition or higher when an aircraft increases power to increase altitude or speed from an approach condition. As used herein, cruise condition includes those abnormal operating conditions that may define pressures, flows, and/or temperatures within the engine 10 generally similar to cruise condition. Still further, although the engine 10 is discussed in regard to within an aircraft, it should be appreciated that similar conditions may be defined in other apparatuses, including, but not limited to, power generation (e.g., industrial, auxiliary power unit, etc.), marine, locomotive, etc.

The first turbine bearing 200 supports the low speed turbine rotor 110 inward of the plurality of outer shroud airfoils 118 at the first stage 101. For example, the first turbine bearing 200 may support the overhung or cantilevered low speed turbine rotor 110 generally forward or upstream 99 of the high speed turbine rotor 120. Still further, the first turbine bearing 200 supports the high speed turbine rotor 120 and the HP shaft 34 coupled to the HP compressor 24.

The arrangement of the first bearing 200 may provide support toward the upstream end 99 of the low speed turbine rotor 110 to be interdigitated forward and/or among the high speed turbine rotor 120. Furthermore, the first bearing 200 provides support toward the upstream end 99 of the low speed turbine rotor 110 that limits an overhanging or cantilevered weight of the low speed turbine rotor 110 from the connecting airfoil 116 upstream toward the combustion section 26. Still further, the first bearing 200 provides support toward the upstream end 99 of the low speed turbine rotor 110 that provides balance to the inner shroud 112 and the plurality of inner shroud airfoils 119 extended therefrom toward the downstream end 98 of the turbine section 90. In various embodiments, the axially extended hub 105 of the low speed turbine rotor 110 may further define one or more balance planes. The balance plane may define features to which weight may be added to or removed from the low speed turbine rotor 110 to aid rotor balance and operation.

Referring now to FIG. 6, another exemplary embodiment of the engine 10 including the first turbine bearing 200 supporting the low speed turbine rotor 110 and the high speed turbine rotor 120 is generally provided. The exemplary embodiment provided in FIG. 6 may be configured substantially similarly as embodiments shown and described in regard to FIGS. 1-5. However, the embodiment of the engine 10 in FIG. 6 further includes a fourth manifold 290 extended from the pressure plenum 84 and coupled to the first turbine bearing 200. The fourth manifold 290 provides fluid communication of flow and pressure of air from the pressure plenum 84 to the first turbine bearing 200, such as shown schematically by arrows 291, through the first turbine bearing 200 and egressing through the cooling orifice 230. The fourth manifold 290 may collect air from the pressure plenum 84 independently of the first manifold 260 and the compressed air 261 pressurized therethrough to the outer air bearing 210 and the inner air bearing 220. As such, the fourth manifold 290 may define orifices, volumes, or areas, including, but not limited to, restrictor or orifice plates, variable geometry, valves, etc. that may induce a different pressure of compressed air 291 through first turbine bearing 200 through the cooling orifice 230 and egressing into the turbine cooling conduit 125 of the high speed turbine rotor 120 (as shown schematically by arrows 231) than that of the compressed air 261.

In various embodiments, the fourth manifold 290 may extend to a separate external compressed air source (e.g., outside of the pressure plenum 84 or outside of the combustion section 26) to induce a flow and/or pressure independent of the flow or pressure through the pressure plenum 84. The flow and/or pressure of air 291 through the fourth manifold 290 may be actively or passively controlled separately from the air 261 through the first manifold 260 into the first turbine bearing 200. Active control may include one or more of a valve or actuation to open or restrict air 291 to the turbine cooling circuit 125. Passive control may include one or more of a metering or restrictor plate, orifice, or nozzle, conduit, or manifold defining an area of volume sized to provide a desired minimum or maximum flow and/or pressure at one or more engine operating conditions.

Referring now to FIG. 7, an exemplary flowchart depicting flows of air and combustion gases through the engine 10 is generally provided. FIG. 7 depicts flows as shown and described in regard to FIGS. 1-6, and various embodiments thereof. However, it should be appreciated that the flowchart may include or omit any one or several embodiments as discussed in regard to each of FIGS. 1-6 or embodiments thereof. FIG. 7 shows an initial flow of air 80 entering the core flowpath 70 of the engine 10. The air 80 enters the IP compressor 22 (or, in two-spool embodiments, the LP compressor) and the HP compressor 24 to become compressed air 82 entering the combustion section 26. A portion of the compressed air 82 egresses from the pressure plenum 84 and enters into the first turbine bearing 200, as shown schematically at arrows 261. As shown and discussed in regard to FIGS. 1-5, the compressed air 261 egresses through the outer plurality of orifices 213 of the outer air bearing 210 and the inner plurality of orifices 223 of the inner air bearing 220, each of which generate a gap or clearance in support of the low speed turbine rotor 110 and the high speed turbine rotor 120, respectively.

As shown and discussed in regard to FIGS. 1-6, the compressed air 261, after egressing from the outer air bearing 210 and inner air bearing 220, is collected in one or more cavities, such as the annular cavity 218, the first cavity 245, and/or the second cavity 255 (FIGS. 1-6), and egressed from the first turbine bearing 200, as shown schematically by arrows 271, to the pressure regulating valve 300.

Referring still to FIG. 7, a portion of compressed air 82 may be egressed from the pressure plenum 84 directly to the pressure regulating valve 300, such as shown schematically by arrows 281. The compressed air 281 provides a reference pressure or motive pressure at the pressure regulating valve 300 in comparison to the input pressure from the compressed air 271 egressing the first turbine bearing 200 and a differential from the air 299 egressing the pressure regulating valve 300 to a lower pressure sink (e.g., such as the core flowpath 70 downstream of the turbine section 90, or an ambient condition).

In various embodiments, such as shown in FIGS. 6-7, compressed air 291 may enter the first turbine bearing 200 and egress into high speed turbine rotor 120, as shown schematically by arrows 231, independently of the compressed air 261 at least partially going through the outer air bearing 210 and inner air bearing 220.

In still various embodiments, at least a portion of air may egress from the first turbine bearing 200 to the core flowpath at or upstream of the turbine section 90, as shown schematically by arrows 262.

Referring still to FIG. 7, a portion of air may egress from the pressure regulating valve 300 into the OD secondary flowpath 71 of the turbine section 90, such as shown schematically by arrows 276. The compressed air 276 may provide cooling to the low speed turbine rotor 110 interdigitated among the high speed turbine rotor 120. In various embodiments, the one way valve 310 is disposed upstream of the OD secondary flowpath 71 and parallel to the flow of compressed air 271 going to the pressure regulating valve 300. The one way valve 310 may restrict a flow of the air 276 toward the OD secondary flowpath 71 and disable flow toward the pressure regulating valve 300.

Referring now to FIG. 8, a flowchart outlining an exemplary method of operating a gas turbine engine 800 is generally provided (hereinafter referred to as “method 800”). The method 800 generally provides an operation of an engine such as that shown in FIGS. 1-7 and described in regard thereto. The engine 10 generally includes a first turbine bearing 200 defining an outer air bearing 210 radially adjacent to a low speed turbine rotor 110 and an inner air bearing 210 radially adjacent to a high speed turbine rotor 120 and an HP shaft 34 coupled thereto. The engine 10 includes a first manifold 260 extended to the first turbine bearing 200 and in fluid communication with a pressure plenum 84 of a combustion section 26 of the engine 10. A second manifold 270 is coupled to the first turbine bearing 200. The second manifold 270 provides fluid communication from the first turbine bearing 200 to the pressure regulating valve 300 or to a lower pressure sink (e.g., an ambient or atmospheric condition outside of the engine, or a lower pressure, such as in the core flowpath proximate to the downstream end of the engine 10). The method 800 outlined in FIG. 8 and described herein may provide operation and support of the interdigitated turbine section 90, such as of the low speed turbine rotor 110 and the high speed turbine rotor 120.

The method 800 includes at 810 flowing air from the pressure plenum to the first turbine bearing; at 820 flowing air between the outer air bearing and the low speed turbine rotor hub; at 830 flowing air between the inner air bearing and the HP shaft; and at 840 flowing air from the first turbine bearing to the pressure regulating valve via the second manifold. Although FIG. 8 depicts steps performed in a particular order for the purposes of illustration and discussion, those of ordinary skill in the art, using the disclosures provided herein, will understand that various steps of any of the methods disclosed herein may be adapted, modified, rearranged, omitted, or expanded in various ways without deviating from the scope of the present disclosure.

In various embodiments of the method 800 outlined in FIG. 8 and described herein, in conjunction with the embodiments of the engine 10 shown in FIGS. 1-7, the step at 810 may include flowing air 261 from the pressure plenum 84 to the first turbine bearing 200 including the outer turbine bearing 210 and the inner turbine bearing 220. At 820, flowing air 261 between the outer air bearing 220 and the hub 105 of the low speed turbine rotor 110 includes defining a pressure of the air 261 between the outer air bearing 220 and the hub 105 such that defines a non-contacting gap or clearance between the first turbine bearing 200 and the hub 105 of the low speed turbine rotor 110. In one embodiment at 830, flowing air between the inner air bearing 220 and the HP shaft 34 defines a pressure of the air 261 such that defines a non-contacting gap or clearance between the first turbine bearing 200 and the HP shaft 34 coupled to the high speed turbine rotor 120. In an embodiment at 840, flowing air 261 from the first turbine bearing 200 to the pressure regulating valve 200 via the second manifold 270 defines a flow and pressure of the air 271 as shown and described in regard to FIGS. 3-7.

In one embodiment, the method 800 further includes at 822 providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub to the second manifold; at 824 providing a second supply of air from between the inner air bearing and the HP shaft to the second manifold; and at 825 flowing the first supply of air and the second supply of air through the second manifold to a lower pressure sink. For example, the method 800 at 822 may include providing air 271 from the annular cavity 218 defined by the groove 217 in the outer air bearing 210, and egressing the air 271 through the first turbine bearing 200 to the second manifold 270. The method 800 at 824 may include providing air 271 from the first cavity 245 through the first turbine bearing 200 to the second manifold 270. The method 800 at 824 may further include providing air 271 from the second cavity 255 through the first turbine bearing 200 to the second manifold 270. The method 800 at 825 may include collecting the air 271 from the annular cavity 218 and the first cavity 245 such that the pressure of the air 271 is normalized, and egressed through the second manifold 270, as shown schematically by arrows 271 in FIGS. 3-7. The method 800 at 825 may further include collecting the air 271 from the annular cavity 218, the first cavity 245, and the second cavity 255, normalizing the pressure, and egressing the air 271 through the second manifold 270.

In one embodiment, the step at 825 includes flowing the first supply of air and the second supply of air through the second manifold 270 to the pressure regulating valve 300. In still various embodiments, the step at 825 may include flowing the first supply of air and the second supply of air through the second manifold to a lower pressure sink. For example, the lower pressure sink may be an ambient or atmospheric pressure condition outside of the engine 10 or toward a downstream end 98 of the core flowpath 70 of the turbine section 90. In various examples, the lower pressure sink generally defines a pressure less than within the pressure plenum 84, the first turbine bearing 200, or the second manifold 270. In one embodiment, the lower pressure sink is approximately 1 bar or less.

In one embodiment of the method 800, the method 800 further includes at 850 flowing air from the pressure plenum of the combustion section to the pressure regulating valve via the third manifold. For example, the method 800, such as implemented in conjunction with the engine 10 shown in FIGS. 1-7, may include flowing or communicating to the pressure regulating valve 300 the pressure within the pressure plenum 84 via the third manifold 280, such as shown via the compressed air 281 from the pressure plenum 84 to the pressure regulating valve 300.

In another embodiment, the method 800 further includes at 860 flowing air from between the outer air bearing and the low speed turbine rotor hub to the core flowpath of the engine at the turbine section. For example, referring to FIGS. 1-7, the method 800 at 860 includes directing a flow of air 262 from the outer plurality of orifices 213 of the outer air bearing 210 to the core flowpath 70 at the turbine section 90 of the engine 10.

Referring now to FIG. 9, a flowchart outlining an exemplary method of operating a gas turbine engine 900 is generally provided (hereinafter referred to as “method 900”). The method 900 generally provides an operation of an engine such as that shown in FIGS. 1-7 and described in regard thereto. The method 900 may further include one or more steps outlined and described in regard to method 800 outlined and described in regard to FIG. 8. The engine 10 generally includes a first turbine bearing 200 defining an outer air bearing 210 radially adjacent to a low speed turbine rotor 110 and an inner air bearing 210 radially adjacent to a high speed turbine rotor 120 and an HP shaft 34 coupled thereto. The engine 10 includes a first manifold 260 extended to the first turbine bearing 200 and in fluid communication with a pressure plenum 84 of a combustion section 26 of the engine 10. A second manifold 270 is coupled to the first turbine bearing 200. A third manifold 280 is coupled to the combustion section 26 and provides fluid communication from the pressure plenum 84 to a pressure regulating valve 300. The second manifold 270 provides fluid communication from the first turbine bearing 200 to the pressure regulating valve 300. The method 900 outlined in FIG. 9 and described herein may provide operation and support of the interdigitated turbine section 90, such as of the low speed turbine rotor 110 and the high speed turbine rotor 120, while also providing thermal attenuation at the high speed turbine rotor 120.

The method 900 may include at 910 flowing air from the pressure plenum to the first turbine bearing; at 920 flowing air between the outer air bearing and the low speed turbine rotor hub; at 930 flowing air between the inner air bearing and the HP shaft; at 940 flowing air from the first turbine bearing to the pressure regulating valve via the second manifold; and at 945 flowing air from the first turbine bearing through a turbine cooling circuit of the high speed turbine rotor. The steps at 910, 920, 930, and 940, and various embodiments thereof, may be undertaken as described in regard to the steps at 810, 820, 830, and 840, respectively, and in various embodiments thereof.

In various embodiments, such as in regard to FIGS. 4-6, the step at 945 includes egressing the flow of air 231 from the first turbine bearing 200 through the cooling orifice 230 into the turbine cooling conduit 125 of the high speed turbine rotor 120. For example, in one embodiment, in reference to the engine 10 shown and described in regard to FIGS. 4-5, the first turbine bearing 200 defines a conduit there through to provide compressed air 261 from the first manifold 260 and pressure plenum 84 and egressing through the cooling orifice 230 adjacent to the turbine cooling circuit 125 of the high speed turbine rotor 120, as depicted schematically by arrows 231.

In one embodiment, the method 900 includes at 950 flowing air from the pressure plenum of the combustion section to the pressure regulating valve via the third manifold, such as described in regard to step 850 of method 800 and shown and described in regard to FIGS. 1-7.

In another embodiment, the method 900 includes at 960 flowing air from between the outer air bearing and the low speed turbine rotor hub to the core flowpath of the turbine section, such as described in regard to step 860 of method 800 and shown and described in regard to FIGS. 1-8.

In still another embodiment, flowing air to the turbine cooling circuit 125 of the high speed turbine rotor 120 includes at 970 modulating or adjusting a flow and/or pressure of the compressed air 231 egressing the first turbine bearing 200 into the turbine cooling circuit 125. For example, referring to FIGS. 6-7, modulating or adjusting a flow and/or pressure of compressed air 231 may include flowing air from the fourth manifold 290 independently of the first manifold 260.

In still another embodiment, the method 900 may further include at 980 monitoring a difference in air pressure from the second manifold and air pressure from the third manifold. For example, monitoring the difference in air pressure includes a difference between the compressed air 271 in the second manifold 270 and the compressed air 281 in the third manifold 280. The compressed air 281, or more specifically, the pressure of the compressed air 281 provides to the pressure regulating valve 300 a reference pressure or muscle air for operating the pressure regulating valve 300 to maintain a desired pressure of the compressed air 261 at the first turbine bearing 200 for maintaining a desired gap or clearance between the first turbine bearing 200 and each of the hub 105 and HP shaft 34.

In various embodiments, the method 900 may include at 922 providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub; at 924 providing a second supply of air from between the inner air bearing and the HP shaft; and at 926 flowing the first supply of air and the second supply of air through the second manifold to the pressure regulating valve. The steps at 922 and 924 may be performed substantially similarly as described in regard to steps 822 and 824 of method 800.

The method 900 at 926 may include collecting the air 271 from the annular cavity 218 and the first cavity 245 such that the pressure of the air 271 is normalized, and egressed through the second manifold 270, as shown schematically by arrows 271 in FIGS. 3-7. The method 900 at 926 may further include collecting the air 271 from the annular cavity 218, the first cavity 245, and the second cavity 255, normalizing the pressure, and egressing the air 271 through the second manifold 270. For example, in one embodiment, the step at 926 includes flowing the first supply of air and the second supply of air through the second manifold 270 to the pressure regulating valve 300.

Referring now to FIG. 10, a flowchart outlining an exemplary method of thermal management a gas turbine engine 1000 is generally provided (hereinafter referred to as “method 1000”). The method 1000 may provide thermal attenuation of the low speed turbine rotor 110 of the interdigitated turbine section 90 such as shown and described in regard to FIGS. 1-7. The method 1000 may further include one or more steps of the method 800 and/or method 900 and outlined and described in regard to FIGS. 8-9.

The method 1000 may include at 1010 flowing air from the pressure plenum to the first turbine bearing; at 1020 flowing air between the outer air bearing and the low speed turbine rotor hub; at 1030 flowing air between the inner air bearing and the HP shaft; and at 1040 flowing air from the first turbine bearing to the pressure regulating valve via the second manifold. The steps at 1010, 1020, 1030, and 1040, and various embodiments thereof, may be undertaken as described in regard to the steps at 810, 820, 830, and 840, respectively, and in various embodiments thereof

The method 1000 further includes at 1047 flowing air from the first turbine bearing to the outer diameter (OD) secondary flowpath via the second manifold. For example, the method 1000 at 1047 may include flowing compressed air 271 from the second manifold 270 through the turbine cooling supply manifold 275 coupled to the outer turbine casing 155 and in fluid communication with the OD secondary flowpath 71, such as shown schematically by arrows 276 in FIG. 5.

In one embodiment of the method 1000 at 1047, flowing air 276 through the OD secondary flowpath 71 may include opening a one way valve (e.g., one way valve 310), such as by defining a pressure of the air 271 in the second manifold 270 that overcomes the one way valve 310 at the turbine cooling supply manifold 275. In various embodiments, the method 1000 at 1047 may include operating the engine 10 at a cruise power condition or higher of the LTO.

In one embodiment, the method 1000 may further includes at 1049 flowing air from the OD secondary flowpath to the core flowpath. For example, in regard to FIG. 5, compressed air, shown schematically by arrows 276, flows from the turbine cooling supply manifold 275 through the OD secondary flowpath 71 into the core flowpath 70. More specifically, the compressed air 276 flows proximate to the outer shroud 114 and the outer shroud airfoils 118 of the low speed turbine rotor 110 to provide cooling. In various embodiments, the compressed air 276 may further flow toward the downstream end 98 within the OD secondary flowpath 71 to further provide cooling along the low speed turbine 110.

In another embodiment, the method 1000 may further include at 1090 flowing air from the pressure regulating valve to a lower pressure sink. For example, referring to FIG. 5, the method 1000 at 1090 may include flowing air 299 from the pressure regulating valve 300 to an active clearance control (ACC) system to govern or control a radial clearance or gap between the outer shroud 114 of the low speed turbine rotor 110 and the plurality of high speed turbine airfoils 122 of the high speed turbine 120, or to an outer diameter seal assembly (e.g., seal assembly 399, 400), and to the core flowpath 70 at the downstream end 98 of the turbine section 90 (e.g., proximate to the inner shroud airfoils 119 of the low speed turbine rotor 110). As another example, the method 1000 at 1090 may include flowing air 299 to an ambient condition outside of the core flowpath 70. In various embodiments, the method 1000 at 1090 includes operating the engine 10 below a cruise power condition (e.g., approach, landing, taxi, or generally at low pressure/low altitude conditions).

In various embodiments, the method 1000 may further includes steps shown and described in regard to method 800 and method 900, such as, the steps 822, 824, 926, 945, 850, 860, 970, or 980 commensurate to steps 1022, 1024, 1026, 1045, 1050, 1060, 1070, or 1080 herein.

Referring now to FIG. 11, another exemplary embodiment of a method of thermal management or active clearance control of an interdigitated turbine section is generally provided (hereinafter, “method 1100”). The method 1100 may further include one or more steps of the methods 800, 900, and/or 1000 and outlined and described in regard to FIGS. 8-10. The method 1100 provides control of clearance gaps between the outer shroud 114 of the low speed turbine rotor 110 and radially outward end or blade tips of the plurality of high speed turbine airfoils 122 interdigitated among the plurality of outer shroud airfoils 118 and radially inward of the outer shroud 114. The method 1100 generally mitigates leakage, or excessive leakage, between the rotating components of the turbine section 90, as well as between the rotating and static components, such as between one or more turbine frames including an aspirating face seal assembly between an outer diameter of the outer shroud 114 and the frame 295.

One embodiment of the method 1100 includes at 1110 flowing air from a compressor section to a rotatable outer shroud of an interdigitated turbine section; and at 1120 adjusting the flow of air to the outer shroud based at least on an engine condition of the turbine engine. In various embodiments, flowing air from the compressor section may include flowing air directly from the compressor section 21 to the turbine section 90. Other embodiments include flowing air such as described in regard to methods 800, 900, and 1000, such as through the first turbine bearing 200.

In one embodiment at 1120, adjusting the flow of air to the outer shroud is based at least on an exhaust gas temperature (EGT) at the turbine section and a rotational speed of the outer shroud. For example, adjusting the flow of air is based on a function, lookup table, transfer function, or curve relating thermal expansion coefficients of the rotary members of the turbine section 90 (e.g., the low speed turbine rotor 110 generally, or the outer shroud 114 specifically, the high speed turbine rotor 120, the plurality of high speed turbine rotor airfoils 122, etc.). Adjusting the flow of air may be further based on one or more of a rotational speed of the low speed turbine rotor 110, the high speed turbine rotor 120, or portions thereof (e.g., the outer shroud 114, the outer shroud airfoils 118, the second turbine airfoils 122, etc.). As such, in various embodiments at 1120, adjusting the flow of air is further based on a rotational speed of a second turbine rotor interdigitated with a first turbine rotor including the outer shroud.

In another embodiment, the method 1100 at 1120 further includes at 1130 determining a desired clearance gap between the outer shroud and a radially outward end of a plurality of second turbine airfoils of the second turbine rotor. The desired clearance gap is generally a function of a desired or maximum acceptable leakage of air across one or more seals (e.g., an aspirating face seal assembly), or between the radially outward end or tip of the plurality of second turbine airfoils 122 and the inner diameter of the outer shroud 114 of the low speed turbine rotor 110.

As another embodiment at 1120, the method 1100 further includes at 1140 determining a desired radial gap between an outer diameter seal assembly and the outer shroud; and at 1150 determining a desired axial gap between the outer diameter seal assembly and the outer shroud. In still another embodiment, the method 1100 further includes at 1160 pressurizing a first side of an outer diameter seal assembly, wherein pressurizing the first side provides a flow of air between the outer shroud and the outer diameter seal assembly. For example, pressurizing the first side of the outer diameter seal assembly comprises pressurizing an aspirating face seal assembly between the outer shroud and a turbine frame.

Referring briefly to FIG. 12, a schematic cross sectional view of another exemplary embodiment of the turbine section 90 is generally provided. The turbine section 90 is configured substantially similarly as shown and described in regard to FIG. 2. In the embodiment generally provided in FIG. 12, the turbine frame 295 further defines a generally longitudinally extended static arm 296. The static arm 296 is extended between the outer shroud 114 of the low speed turbine rotor 110 and an outer wall 297 of the turbine frame 295. The outer shroud 114 further defines a generally longitudinally extended rotary arm 117 extended between the turbine frame 295 and the outer shroud 114. The rotary arm 117 and the static arm 296 each extend generally annularly around the axial centerline 12 radially between the outer diameter 297 of the turbine frame 295 and the outer shroud 114.

In one embodiment, the longitudinally extended rotary arm 117 and static arm 296 together define the labyrinth seal assembly 399 including a shroud or grooves at the static arm 296 and a plurality of teeth at the rotary arm 117 extended to the static arm 296. In other embodiments, the longitudinally extended rotary arm 117 and static arm 296 may together define the outer diameter seal assembly 400 defining the aspirating face seal assembly, such as shown and described in regard to FIGS. 13-15.

Referring now to FIG. 13, the outer diameter seal assembly 400 may define the aspirating face seal assembly between the turbine frame 295 and the outer shroud 114 of the low speed turbine 110 generally provided. The outer diameter seal assembly 400 defines a primary tooth 401 radially inward of a longitudinal face 350 of the turbine frame 295. The space between primary tooth 401 and the longitudinal face 350 of the turbine frame 295 defines a primary tooth gap 315. The outer diameter seal assembly 400 further defines a secondary tooth 402 axially adjacent to a radial face 360 of the turbine frame 295. The space between the secondary tooth 402 and the radial face 360 of the turbine frame 295 defines a secondary tooth gap 325. At relatively low power or low temperatures at the turbine section 90 (e.g., startup, idle, flight idle, etc.), the primary tooth gap 315 is greater than the secondary tooth gap 325 such as to permit flow of air therethrough. As operating conditions increase, the secondary or starter tooth gap 325 increases greater than the primary tooth gap 315.

As further provided in FIG. 13, a flow of air, shown schematically by arrows 410, enters and egresses through a plurality of feed holes 370 through the radial face 360 of the turbine frame 295. At a desired operating condition (e.g., the aero design point, or one or more of mid-power and high-power conditions, including, but not limited to, cruise, climb, takeoff, or approach), the primary tooth 401 acts as the main flow restriction to mitigate undesired leakage across the outer diameter seal assembly 400, thereby improving efficiency of the interdigitated turbine section 90.

Another flow of air, shown schematically by arrows 420, enters and egresses across the radial gap 315 and the axial gap 325. The flows 410 and 420 merge between the turbine frame 295 and the outer shroud 114 to become a vent flow, shown schematically at 430. The vent flow 420 egresses through a plurality of vent openings 425 through the turbine frame 295 and between the turbine frame 295 and the outer shroud 114 to egress in fluid communication with the relatively lower pressure core flowpath 70.

The plurality of vent openings 425 may generally define discrete passages extended through the turbine frame 295 (e.g., orifices or holes through the turbine frame 295 extended therethrough as conduits or passages). The plurality of vent openings 425 may be disposed in adjacent circumferential arrangement though the turbine frame 295. It should be appreciated that the vent openings 425 may define substantially circular cross sections, or ovular, oblong, rectangular, or polygonal cross sectional areas. Still further, the vent openings 425 may define varying cross sectional areas therethrough, such as to increase or decrease a flow or pressure of the air egressing therethrough.

In still various embodiments, the plurality of feed holes 370 may generally define discrete passages extended through the turbine frame 295 such as described in regard to the vent openings 425. Each of the pluralities of feed holes 370 and vent openings 425 may together be in alternating circumferential arrangement such as to be fluidly disconnected from one another (e.g., the cross section shown in FIG. 13 is depicted at a first circumferential location at the feed hole 370 and at a second circumferential location at the vent opening 425).

Referring now to FIG. 14, another cross sectional view of a portion of the turbine section 90 including the seal assembly 400 at the turbine frame 295 and the outer shroud 114 is generally provided. The seal assembly 400 generally includes a plurality of feed holes 370 disposed in adjacent circumferential arrangement around the axial centerline 12 through the turbine frame 295. The feed holes 370 are in fluid communication with a manifold providing air from the compressor section 21.

In one embodiment, the manifold is a fifth manifold 292 extended from the pressure regulating valve 300 (shown in FIGS. 3-7) to the turbine frame 295 and coupled thereto, providing a pressurized flow of air 299, 410 to the seal assembly 399, 400. In another embodiment, the manifold 292 is coupled directly from the compressor section 21 (e.g., the low pressure compressor 22 or the high pressure compressor 24) to the turbine frame 295 to provide the flow of air 299 to the seal assembly 399, 400, such as to provide the air 299, 410, 420 as described above.

Referring now to FIG. 15, the cross sectional view of the turbine section 90 including the seal assembly 400 generally provided at least partially includes a schematic representation of exemplary forces or loads acting upon and within the seal assembly 400, the outer shroud 114, and the turbine frame 295. The turbine frame 295 may include a generally elastic or springing structure 500 to provide displacement of the static structure, such as the turbine frame 295. In various embodiments, the spring 500 produces a reacting force, shown schematically by arrows 510, in response to a high pressure closing force, shown schematically by arrows 515, of the turbine frame 295 relative to the outer shroud 114 along the longitudinal direction L. For example, the loading exerted onto the radial face 350 of the turbine frame 295 toward the radial portion 380 of the outer shroud 114 downstream of and adjacent to the radial face 350 of the turbine frame 295 is reacted against the spring 500.

The aspirating seal assembly 400 generally includes at the radial face 350 of the turbine frame 295 a low pressure closing force, shown schematically by arrows 520, exerted toward the radial portion 380 of the outer shroud 114 within a cavity 550 defined between the turbine frame 295 and a longitudinal portion 115 of the outer shroud 114. For example, the low pressure closing force 520 is generally co-directional to the high pressure closing force 515.

The aspirating seal assembly 400 further includes an opening force, shown schematically by arrows 530, adjacent to the secondary tooth 402 extended generally along the longitudinal direction proximate to the radial face 350 of the turbine frame 295. The opening force 530 is generally exerted by the high pressure air 410 entering the seal assembly 400 from the manifold 292.

Still further, the aspirating seal assembly 400 further defines an air bearing, shown schematically by arrows 525, between the radial face 350 of the turbine frame 295 and the radial portion 380 of the outer shroud 114. The air bearing 525 is defined, at least in part, by the high pressure flow 410 across the primary tooth 401 and the secondary tooth 402. Still further, the air bearing 525 is defined at least in part by the high pressure flow 420 entering through the feed holes 370. As such, the air bearing 525 provides a layer of fluid between the rotating outer shroud 114 and the static turbine frame 295.

The various embodiments of the aspirating seal assembly 400 generally provided include an approximately annular aspirating seal assembly 400 extended around the circumference of the outer shroud 114 and turbine frame 295. The primary tooth 401 and the secondary tooth 402 may each extend substantially or entirely circumferentially and generally proximate to an annular turbine frame 295. As such, the primary tooth 401 and the secondary tooth 402 may together provide sealing of air between the rotating outer shroud 114 and the static turbine frame 295, thereby improving engine efficiency and performance, and mitigating leakages that may adversely affect the interdigitated turbine section 90.

As previously noted, the seal assemblies 399, 400 may receive a portion of air directly from the compressor section 21, such as to provide the flows 410 and 420 described herein. In other embodiments, the seal assembly 400 may receive a portion of air from the first turbine bearing assembly 200, such as generally provided by flow 299 from the pressure regulating valve 300. In still other embodiments, the seal assembly 400 may receive a portion of air from the compressor section 21 to a pressure regulating valve 300 such as to provide a generally uniform pressure around the circumference of the outer shroud 114 and seal assembly 399, 400.

Referring back to FIG. 11 and method 1100, in various embodiments, determining the desired radial gap at 1140 includes determining the desired distance or displacement at the primary tooth gap 315 between the longitudinal face 350 of the turbine frame 295 and the primary tooth 401. In other embodiments, determining the desired axial gap at 1150 includes determining the desired distance or displacement at the secondary tooth gap 325 between the radial face 360 and the secondary tooth 402.

Referring to another embodiment, the method 1100 further includes at 1170 flowing air from the compressor section to a second turbine rotor (e.g., high speed turbine rotor 120) interdigitated with a plurality of outer shroud airfoils of the outer shroud (e.g., outer shroud airfoils 118 extended radially inward from the outer shroud 114). In one embodiment, flowing air to the second turbine rotor includes flowing air to an inner diameter of the second turbine rotor to attenuate radial thermal growth. For example, referring back to FIGS. 3-7, the inner diameter may include the turbine cooling circuit 125 of the high speed turbine rotor 120. The turbine cooling circuit 125 may dispose the flow of air to the plurality of high speed turbine rotor airfoils 122 to help mitigate or control thermal growth and thereby control a gap or clearance (e.g., a radial gap or clearance between the plurality of high speed turbine airfoils 122 and an inner diameter of the outer shroud 114, such as adjacent within the core flowpath 70).

In another embodiment, the method 1100 includes at 1172 generating a flow and pressure of air through a primary flowpath (e.g., core flowpath 70) and a secondary flowpath (e.g., secondary flowpath 71); at 1174 directing at least a portion of the air to a seal assembly (e.g., labyrinth seal assembly 399, aspirating seal assembly 400) between the outer shroud and a turbine frame (e.g., outer shroud 114 and turbine frame 295); and at 1176 directing at least a portion of the air along a longitudinal direction of and around a circumference of the outer shroud based on one or more engine operating conditions (e.g., the landing/takeoff cycle).

In still another embodiment, the method 1100 further includes at 1178 directing at least a portion of air to the second turbine rotor disposed radially inward of the outer shroud. The portion of air provides thermal attenuation at the second turbine rotor to mitigate radial growth of the second turbine rotor.

In still yet another embodiment, the method 1100 includes at 1180 determining a desired flow or pressure of air at one or more of the outer shroud and second turbine rotor based on one or more of a lookup table, a transfer function, and a curve. As previously mentioned, the function, curve, table, etc. is at least a function of radial dimensions of the outer shroud 114 and the second turbine rotor (e.g., high speed turbine rotor 120) relative to one or more of a rotational speed of the outer shroud 114, a rotational speed of the second turbine rotor (e.g., high speed turbine rotor 120), an exhaust gas temperature (e.g., temperature of the combustion gases 86 at the turbine section 90), a material thermal expansion constant or coefficient relative to the outer shroud 114 of the low speed turbine rotor 110, and a material thermal expansion constant or coefficient relative to the second turbine rotor (e.g., high speed turbine rotor 120, or more specifically, the plurality of high speed turbine rotor airfoils 122).

The turbine section 90 shown and described herein may improve upon existing turbine sections by providing improved fuel efficiency, operational efficiency, and/or power output while maintaining or reducing weight, part count, and/or packaging. The methods of clearance control and structures mitigating leakage at the outer diameter rotor to static interface (e.g., the outer shroud 114 to turbine frame 295) may enable performance and efficiency improvements to the engine 10 via the interdigitated turbine section 90 while mitigating losses related to outer diameter leakages. The plurality of outer shroud airfoils 118 of the low speed turbine rotor 110 interdigitated among the plurality of high speed turbine airfoils 122 of the high speed turbine rotor 120 may reduce packaging and reduce part count by removing stages of stationary airfoils between each rotating component. Additionally, the turbine section 90 may provide efficiency benefits comparable to a reduction gearbox without adding weight or size (e.g. axial length) to the engine 10. The low speed turbine rotor 110, as the first stage downstream of the combustion section 26, may further improve engine efficiency by reducing cooling air appropriated away from producing combustion gases 86, thereby allowing more energy from the compressor section 21 to be used in combustion and operation of the engine 10. Furthermore, removing the nozzle guide vane between the combustion section 26 and the low speed turbine rotor 110 of the turbine section 90 may reduce or eliminate design constraints related to hot spots in the combustion gases along the annulus of the core flowpath 70.

The various embodiments of the turbine section 90 generally shown and described herein may be constructed as individual blades installed into drums or hubs, or integrally bladed rotors (IBRs) or bladed disks, or combinations thereof. The blades, hubs, or bladed disks may be formed of ceramic matrix composite (CMC) materials and/or metals appropriate for gas turbine engine hot sections, such as, but not limited to, nickel-based alloys, cobalt-based alloys, iron-based alloys, or titanium-based alloys, each of which may include, but are not limited to, chromium, cobalt, tungsten, tantalum, molybdenum, and/or rhenium. The turbine section 90, or portions or combinations of portions thereof, including the inner shroud 112, the outer shroud 114, the connecting airfoil(s) 116, the plurality of outer shroud airfoils 118, and/or the plurality of inner shroud airfoils 119, may be formed using additive manufacturing or 3D printing, or casting, forging, machining, or castings formed of 3D printed molds, or combinations thereof. The turbine section 90, or portions thereof, such as stages of the rotating components 110, 120, 130, the outer shroud 114, the inner shroud 112, and/or various shrouds, seals, and other details may be mechanically joined using fasteners, such as nuts, bolts, screws, pins, or rivets, or using joining methods, such as welding, brazing, bonding, friction or diffusion bonding, etc., or combinations of fasteners and/or joining methods. Still further, it should be understood that the low speed turbine rotor 110, including the inner and/or outer shroud 112, 114, may incorporate features that allow for differential expansion. Such features include, but are not limited to, aforementioned methods of manufacture, various shrouds, seals, materials, and/or combinations thereof.

The systems and methods shown in FIGS. 1-15 and described herein provide methods and structures for reducing flowpath leakages in outer diameter rotors of the turbine section of an interdigitated turbine engine that may decrease fuel consumption, increase operability, increase engine performance and/or power output while maintaining or reducing weight, part count, and/or packaging (e.g. radial and/or axial dimensions). The systems provided herein may allow for increased high bypass ratios and/or overall pressure ratios over existing gas turbine engine configurations, such as turbofans, while maintaining or reducing packaging relative to other gas turbine engines of similar power output. The systems described herein may contribute to improved bypass ratio and/or overall pressure ratio and thereby increase overall gas turbine engine efficiency.

Still further, the systems and methods shown in FIGS. 1-15 and described herein may reduce a product of a flow area and the square of the rotational speed (the product herein referred to as “AN²”) of the gas turbine engine. For example, engine 10 shown and described in regard to FIGS. 1-15 may generally reduce AN² relative to a conventional geared turbofan configuration. Generally, lowering the AN², such as by reducing the rotational speed and/or the flow area, increases the required average stage work factor (i.e. the average required loading on each stage of rotating airfoils). However, the systems described herein may lower the AN² while also lowering the average stage work factor and maintaining axial length of the turbine section 90 (compared to engines of similar thrust output and packaging) by interdigitating the low speed turbine rotor 110 among the one or more stages of the high speed turbine rotor 120 and the intermediate speed turbine 130 while also defining a non-digitated turbine structure (i.e. the inner shroud 112 and the plurality of inner shroud airfoils 119) toward the downstream end 98 of the turbine section 90. Therefore, the low speed turbine rotor 110 may increase the quantity of rotating stages of airfoils while reducing the average stage work factor, and therefore the AN², while mitigating increases in axial length to produce a similar AN² value. The low speed turbine rotor 110 may further reduce the AN² while additionally reducing the overall quantity of airfoils, rotating and stationary, in the turbine section 90 relative to turbine sections of gas turbine engines of similar power output and/or packaging.

Furthermore, the systems and methods shown in FIGS. 1-15 and described herein may further improve engine efficiency, reduce airfoil quantity, reduce engine weight, and/or alleviate combustion section design constraints by interdigitating the low speed turbine rotor 110 forward or upstream 99 of the high speed turbine rotor 120. For example, defining the first stage of the low speed turbine rotor 110 as immediately downstream 98 of the combustion section 26, without a first turbine vane or nozzle guide vane therebetween, as well as defining the low speed turbine rotor 110 in counter-rotation with the high speed turbine rotor 120, may reduce effects of overall combustion hot spots on the first stage of the low speed turbine rotor 110 in contrast to a stationary, first turbine vane or nozzle guide vane. As such, the turbine section 90 and engine 10 described herein may remove constraints to combustion section 26 design by de-emphasizing hot spots, or combustion pattern factor, in favor of other design criteria, such as decreasing emissions, improving lean blow-out (LBO) and/or altitude re-light, improving overall operability across part or all of an operating envelope, or increasing the operating envelope.

This written description uses examples to disclose the invention, including the best mode, and also to enable any person skilled in the art to practice the invention, including making and using any devices or systems and performing any incorporated methods. The patentable scope of the invention is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they include structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims. 

What is claimed is:
 1. A method of operating an active clearance control system for an interdigitated turbine engine, the method comprising: flowing air from a compressor section to a rotatable outer shroud of an interdigitated turbine section; and adjusting the flow of air to the outer shroud based at least on an engine condition of the turbine engine.
 2. The method of claim 1, wherein adjusting the flow of air to the outer shroud is based at least on an exhaust gas temperature at the turbine section and a rotational speed of the outer shroud.
 3. The method of claim 2, wherein adjusting the flow of air is further based on a rotational speed of a second turbine rotor interdigitated with a first turbine rotor including the outer shroud.
 4. The method of claim 3, wherein adjusting the flow of air to the outer shroud further comprises: determining a desired clearance gap between the outer shroud and a radially outward end of a plurality of second turbine airfoils of the second turbine rotor.
 5. The method of claim 3, wherein adjusting the flow of air to the outer shroud further comprises: determining a desired radial gap between an outer diameter seal assembly and the outer shroud; and determining a desired axial gap between the outer diameter seal assembly and the outer shroud.
 6. The method of claim 3, wherein adjusting the flow of air to the outer shroud further comprises: pressurizing a first side of an outer diameter seal assembly, wherein pressurizing the first side provides a flow of air between the outer shroud and the outer diameter seal assembly.
 7. The method of claim 1, further comprising: flowing air from the compressor section to a second turbine rotor interdigitated with a plurality of outer shroud airfoils of the outer shroud.
 8. The method of claim 7, wherein flowing air to the second turbine rotor includes flowing air to an inner diameter of the second turbine rotor to attenuate radial thermal growth.
 9. The method of claim 6, wherein pressurizing the first side of the outer diameter seal assembly comprises pressurizing an aspirating face seal assembly between the outer shroud and a turbine frame.
 10. The method of claim 1, further comprising: rotating the gas turbine engine to provide a flow and pressure of air through a primary flowpath and a secondary flowpath; directing at least a portion of the air to an aspirating face seal assembly between the outer shroud and a turbine frame; and directing at least a portion of the air along a longitudinal direction of and around a circumference of the outer shroud based on one or more engine operating conditions.
 11. The method of claim 10, further comprising: directing at least a portion of air to the second turbine rotor disposed radially inward of the outer shroud, wherein the portion of air provides thermal attenuation at the second turbine rotor to mitigate radial growth of the second turbine rotor.
 12. The method of claim 1, further comprising: determining a desired flow or pressure of air at one or more of the outer shroud and second turbine rotor based on one or more of a lookup table, a transfer function, and a curve comprising at least a function of radial dimensions of the outer shroud and the second turbine rotor relative to one or more of a rotational speed of the outer shroud, a rotational speed of the second turbine rotor, an exhaust gas temperature, a material thermal expansion constant relative to the outer shroud, and a material thermal expansion constant relative to the second turbine rotor.
 13. The method of claim 1, further comprising: flowing air from the pressure plenum to a first turbine bearing; flowing air between an outer air bearing of the first turbine bearing and a low speed turbine rotor hub; flowing air between an inner air bearing of the first turbine bearing and a high pressure shaft; and flowing air from the first turbine bearing to a pressure regulating valve.
 14. The method of claim 13, further comprising: providing a first supply of air from between the outer air bearing and the low speed turbine rotor hub to a second manifold; and providing a second supply of air from between the inner air bearing and the HP shaft to the second manifold.
 15. A gas turbine engine, comprising: a low speed turbine rotor comprising an inner shroud, an outer shroud, and at least one connecting airfoil coupling the inner shroud and the outer shroud, wherein the outer shroud comprises a plurality of outer shroud airfoils extended inward along a radial direction; a turbine frame at least partially surrounding the low speed turbine rotor; a seal assembly disposed between the outer shroud of the low speed turbine rotor and the turbine frame; and a fifth manifold coupled to the turbine frame, wherein the fifth manifold provides a pressurized flow of air to the seal assembly.
 16. The gas turbine engine of claim 15, wherein the turbine frame defines a longitudinally extended static arm extended between the outer shroud of the low speed turbine rotor and an outer diameter of the turbine frame, and wherein the static arm at least partially defines the seal assembly.
 17. The gas turbine engine of claim 15, wherein the outer shroud of the low speed turbine rotor defines a longitudinally extended rotary arm extended between the turbine frame and the outer shroud, and wherein the rotary arm at least partially defines the seal assembly.
 18. The gas turbine engine of claim 15, wherein the seal assembly defines a labyrinth seal assembly at the turbine frame and the outer shroud of the low speed turbine rotor.
 19. The gas turbine engine of claim 15, wherein the fifth manifold is coupled to a pressure regulating valve in fluid communication with a second manifold providing pressurized air from a first turbine bearing.
 20. The gas turbine engine of claim 15, further comprising: a compressor section, wherein the fifth manifold is coupled to the compressor section and extended to the turbine frame to provide a pressurized flow of air to the seal assembly. 